Panel burn through tolerant shell design

ABSTRACT

A dual wall liner for a gas turbine engine may comprise a shell having a first side and a second side, a panel contacting the shell, the panel at least partially defining a hot gas path through which a hot gas flows, wherein the first side of the shell faces the panel, wherein the shell includes a thermal barrier coating (TBC) disposed on the first side of the shell. The TBC may thermally protect the shell from heat from a hot gas path.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims priority to, and the benefit of, U.S.Provisional Patent Application Ser. No. 62/618,841, entitled “PANEL BURNTHROUGH TOLERANT SHELL DESIGN,” filed on Jan. 18, 2018, The '841application is hereby incorporated by reference in its entirety for allpurposes.

FIELD

The present disclosure relates to protective coatings for turbinecombustion components, and specifically to protective coatings for innersurfaces of turbine combustion components.

BACKGROUND

The efficiency of turbine engines, for example gas turbines, isincreased as the firing temperature, otherwise known as the workingtemperature, of the turbine increases. This increase in temperatureresults in at least some increase in power with the use of the same, ifnot less, fuel. Thus it is desirable to raise the firing temperature ofa turbine to increase the efficiency.

However, as the firing temperature of gas turbines rises, the metaltemperature of the combustion components, including but not limited tocombustion shells, combustion panels, transition pieces otherwise knownas ducts, and augmenters increases. A combustion panel is incorporatedinto a turbine, and defines, in part with a transition piece or duct, anarea for a flame to burn fuel. These components, as well as othercomponents in the gas path environment, such as the combustion shell forexample, are subject to significant temperature extremes and degradationby oxidizing and corrosive environments.

Various techniques have been devised to maintain the temperature of gasturbine components below undesirable levels. For example, a coolantmedium such as coolant air from the compressor of the turbine is oftendirected to a component along one or more component surfaces. Such flowis understood in the art as backside flow where the cooling medium isdirected at a surface of the component not directly exposed to hightemperatures such as the hot gases of combustion. One such component ofthe gas turbine is the combustor panel. It will be appreciated that thecombustor panel confines the hot gases of combustion for flow along thecombustor to a transition body for flow into the turbine section of thegas turbine. The combustor panel may be disposed inward from thecombustor shell, or liner. An annulus may be disposed around the shellwhich receives the coolant air flow on the coolant side of the shell.The metal surface of the combustor shell facing the annulus is normallysmooth.

SUMMARY

A dual wall liner for a gas turbine engine is disclosed, comprising ashell having a first side and a second side, a panel contacting theshell, the panel at least partially defining a hot gas path throughwhich a hot gas flows, wherein the first side of the shell faces thepanel, wherein the shell includes a thermal barrier coating (TBC)disposed on the first side of the shell.

In various embodiments, the panel is coupled to the shell.

In various embodiments, the dual wall liner further comprises a metalliclayer in contact with the TBC.

In various embodiments, the shell is comprised of a shell material, anoxidation resistance of the metallic layer is greater than that of theshell material, and the metallic layer is at least twice as resistant tooxidation as the shell material.

In various embodiments, the TBC is disposed over the metallic layer.

In various embodiments, the TBC comprises a lower thermal conductivitythan that of the shell material.

In various embodiments, the thermal conductivity of the TBC is less thanhalf of that of the shell material.

In various embodiments, the dual wall liner further comprises a heattransfer augmentation feature disposed on the second side of the shell.

In various embodiments, the heat transfer augmentation feature isconfigured to at least one of increase a surface area of the second sideor disturb a flow of cooling air flowing across the second side.

In various embodiments, the dual wall liner further comprises a panelstud coupling the panel to the shell, wherein the TBC surrounds thepanel stud.

In various embodiments, the dual wall liner further comprises aplurality of holes extending through the shell and the TBC, wherein theplurality of holes are configured to direct a flow of cooling air fromthe second side of the shell to the panel.

A gas turbine engine combustor is disclosed, comprising a shell having afirst side and a second side, a panel contacting the shell, the panel atleast partially defining a hot gas path through which a hot gas flows,wherein the first side of the shell faces the panel, wherein the shellincludes a thermal barrier coating (TBC) disposed on the first side ofthe shell.

In various embodiments, the panel is coupled to the shell.

In various embodiments, the gas turbine engine combustor furthercomprises a metallic layer in contact with the TBC.

In various embodiments, the shell is comprised of a shell material, themetallic layer is at least twice as resistant to oxidation as the shellmaterial, the TBC comprises a lower thermal conductivity than that ofthe shell material, and the TBC is disposed over the metallic layer.

In various embodiments, the gas turbine engine combustor furthercomprises a heat transfer augmentation feature disposed on the secondside of the shell, wherein the heat transfer augmentation feature isconfigured to at least one of increase a surface area of the second sideor disturb a flow of cooling air flowing across the second side.

In various embodiments, the gas turbine engine combustor furthercomprises a panel stud coupling the panel to the shell, wherein the TBCsurrounds the panel stud.

In various embodiments, the gas turbine engine combustor furthercomprises a plurality of holes extending through the shell and the TBC,wherein the plurality of holes are configured to direct a flow ofcooling air from the second side of the shell to the panel.

A method for manufacturing a dual wall liner for a gas turbine engine isdisclosed, comprising disposing a thermal barrier coating (TBC) on afirst side of a shell, and coupling a panel to the shell, wherein thefirst side faces the panel.

In various embodiments, the method further comprises disposing ametallic layer on the first side, the TBC disposed over the metalliclayer.

The foregoing features and elements may be combined in variouscombinations without exclusivity, unless expressly indicated hereinotherwise. These features and elements as well as the operation of thedisclosed embodiments will become more apparent in light of thefollowing description and accompanying drawings.

BRIEF DESCRIPTION OF THE DRAWINGS

The subject matter of the present disclosure is particularly pointed outand distinctly claimed in the concluding portion of the specification. Amore complete understanding of the present disclosure, however, may bestbe obtained by referring to the detailed description and claims whenconsidered in connection with the drawing figures, wherein like numeralsdenote like elements.

FIG. 1 is cross-sectional view of an exemplary gas turbine engine inaccordance with various embodiments;

FIG. 2 is cross-sectional view of an exemplary combustor having a dualwall liner, in accordance with various embodiments;

FIG. 3A is cross-sectional view of a dual wall liner having a thermalcoating comprising a metallic layer and a thermal barrier coating (TBC),in accordance with various embodiments;

FIG. 3B is cross-sectional view of the dual wall liner with a panel studcoupling the panel to the shell, in accordance with various embodiments;

FIG. 3C is cross-sectional view of the dual wall liner with a pluralityof holes extending through the shell and the thermal coating and heattransfer augmentation features extending from the shell, in accordancewith various embodiments;

FIG. 4A is cross-sectional view of a dual wall liner having a thermalcoating comprising a metallic layer, in accordance with variousembodiments;

FIG. 4B is cross-sectional view of the dual wall liner with a panel studcoupling the panel to the shell, in accordance with various embodiments;

FIG. 4C is cross-sectional view of the dual wall liner with a pluralityof holes extending through the shell and the thermal coating and heattransfer augmentation features extending from the shell, in accordancewith various embodiments; and

FIG. 5 is a method for manufacturing a dual wall liner, in accordancewith various embodiments.

DETAILED DESCRIPTION

The detailed description of exemplary embodiments herein makes referenceto the accompanying drawings, which show exemplary embodiments by way ofillustration and their best mode. While these exemplary embodiments aredescribed in sufficient detail to enable those skilled in the art topractice the inventions, it should be understood that other embodimentsmay be realized and that logical, chemical and mechanical changes may bemade without departing from the spirit and scope of the inventions.Thus, the detailed description herein is presented for purposes ofillustration only and not of limitation. For example, the steps recitedin any of the method or process descriptions may be executed in anyorder and are not necessarily limited to the order presented.Furthermore, any reference to singular includes plural embodiments, andany reference to more than one component or step may include a singularembodiment or step. Also, any reference to attached, fixed, connected orthe like may include permanent, removable, temporary, partial, fulland/or any other possible attachment option. Additionally, any referenceto without contact (or similar phrases) may also include reduced contactor minimal contact.

As used herein, “aft” refers to the direction associated with the tail(e.g., the back end) of an aircraft, or generally, to the direction ofexhaust of the gas turbine engine. As used herein, “forward” refers tothe direction associated with the nose (e.g., the front end) of anaircraft, or generally, to the direction of flight or motion.

A combustor shell, as disclosed herein, may include a thermal coatingdisposed on an inner surface, also referred to herein as a first side,of the combustor shell. The first side may not be directly exposed tothe hot gas path of the combustor.

Thermal coatings may be costly and may increase the overall weight ofthe engine.

In various embodiments and with reference to FIG. 1, a gas turbineengine 20 is provided. Gas turbine engine 20 may be a two-spool turbofanthat generally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mayinclude, for example, an augmentor section among other systems orfeatures. In operation, fan section 22 can drive air along a bypassflow-path B while compressor section 24 can drive air along a coreflow-path C for compression and communication into combustor section 26then expansion through turbine section 28. Although depicted as aturbofan gas turbine engine 20 herein, it should be understood that theconcepts described herein are not limited to use with turbofans as theteachings may be applied to other types of turbine engines includingthree-spool architectures.

Gas turbine engine 20 may generally comprise a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A-A′ relative to an engine static structure 36 viaseveral bearing systems 38, 38-1, and 38-2. It should be understood thatvarious bearing systems 38 at various locations may alternatively oradditionally be provided, including for example, bearing system 38,bearing system 38-1, and bearing system 38-2.

Low speed spool 30 may generally comprise an inner shaft 40 thatinterconnects a fan 42, a low pressure (or first) compressor section 44and a low pressure (or first) turbine section 46. Inner shaft 40 may beconnected to fan 42 through geared architecture 48 that can drive fan 42at a lower speed than low speed spool 30. Geared architecture 48 maycomprise a gear assembly 60 enclosed within a gear housing 62. Gearassembly 60 couples inner shaft 40 to a rotating fan structure.

High speed spool 32 may comprise an outer shaft 50 that interconnects ahigh pressure (or second) compressor section 52 and high pressure (orsecond) turbine section 54. A combustor 56 may be located between highpressure compressor 52 and high pressure turbine 54. A mid-turbine frame57 of engine static structure 36 may be located generally between highpressure turbine 54 and low pressure turbine 46. Mid-turbine frame 57may support one or more bearing systems 38 in turbine section 28. Innershaft 40 and outer shaft 50 may be concentric and rotate via bearingsystems 38 about the engine central longitudinal axis A-A′, which iscollinear with their longitudinal axes. As used herein, a “highpressure” compressor or turbine experiences a higher pressure than acorresponding “low pressure” compressor or turbine.

The core airflow C may be compressed by low pressure compressor section44 then high pressure compressor 52, mixed and burned with fuel incombustor 56, then expanded over high pressure turbine 54 and lowpressure turbine 46. Mid-turbine frame 57 includes airfoils 59 which arein the core airflow path. Turbines 46, 54 rotationally drive therespective low speed spool 30 and high speed spool 32 in response to theexpansion.

Gas turbine engine 20 may be, for example, a high-bypass geared aircraftengine. In various embodiments, the bypass ratio of gas turbine engine20 may be greater than about six (6). In various embodiments, the bypassratio of gas turbine engine 20 may be greater than ten (10). In variousembodiments, geared architecture 48 may be an epicyclic gear train, suchas a star gear system (sun gear in meshing engagement with a pluralityof star gears supported by a carrier and in meshing engagement with aring gear) or other gear system. Gear architecture 48 may have a gearreduction ratio of greater than about 2.3 and low pressure turbine 46may have a pressure ratio that is greater than about 5. In variousembodiments, the bypass ratio of gas turbine engine 20 is greater thanabout ten (10:1). In various embodiments, the diameter of fan 42 may besignificantly larger than that of the low pressure compressor section44, and the low pressure turbine 46 may have a pressure ratio that isgreater than about 5:1. Low pressure turbine 46 pressure ratio may bemeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of low pressure turbine 46 prior to an exhaustnozzle. It should be understood, however, that the above parameters areexemplary of various embodiments of a suitable geared architectureengine and that the present disclosure contemplates other turbineengines including direct drive turbofans. A gas turbine engine maycomprise an industrial gas turbine (IGT) or a geared engine, such as ageared turbofan, or non-geared engine, such as a turbofan, or maycomprise any gas turbine engine as desired.

With reference to FIG. 2, an exemplary combustor 56 positioned betweencompressor 52 and turbine section 54 of a gas turbine engine 20 having acentral longitudinal axis or centerline axis 100 is illustrated, inaccordance with various embodiments. Combustor 56 may have an annularcombustion chamber 34 defined by inner wall 70 and outer wall 72 andbulkhead 58 spanning the walls 70, 72. Each wall 70, 72 can have one ormore liners 74, 76, 78, 80, with each liner 74, 76, 78, 80 including ashell 68 and panels 75, 77, 79, 81. The panels 75, 77, 79, 81 face thecombustion chamber 34 and shields the shells 68 from the combustionchamber 34. Alternate materials include ceramics and ceramic matrixcomposites. Various known or hereinafter developed other materials andmanufacturing techniques may be utilized. In known fashion or otherwise,the panels may be spaced apart from the interior surface of theassociated shell. The exemplary shells and panels may be foraminate,passing cooling air from annular chambers 90 and 92 respectively throughinboard and outboard of the walls 70 and 72 into the combustion chamber34. The exemplary panels may be configured so that the intact portionsof their inboard surfaces are substantially frustoconical. Viewed inlongitudinal section, these surfaces appear as straight lines atassociated angles to the centerline axis 100. A combusting mixture isdriven downstream within the combustor 56 to a combustor outlet 61immediately ahead of a fixed first vane stage 63 of high pressureturbine 54.

With reference to FIG. 3A, a cross-section view of a dual wall liner 300for a gas turbine engine is illustrated, in accordance with variousembodiments. Dual wall liner 300 may include a shell 310 and anassociated panel 320, in accordance with various embodiments. Shell 310may be similar to shell 68, with momentary reference to FIG. 2. Invarious embodiments, shell 310 may comprise a shell material 311. Invarious embodiments, shell material 311 may comprise a nickel alloy.Panel 320 may be similar to panel 75, panel 77, panel 79, and/or panel81, with momentary reference to FIG. 2. Shell 310 may include a firstsurface 312 and a second surface 314. First surface 312 may face towardspanel 320.

Dual wall liner 300 may further include a thermal coating 330 disposedon first surface 312 of shell 310. In this regard, thermal coating 330may be disposed between shell 310 and panel 320. Thermal coating 330 mayinclude a first layer (also referred to herein as a metallic layer) 332.Metallic layer 332 may be disposed onto first surface 312. Thermalcoating 330 may include a second layer (also referred to herein as athermal barrier coating (TBC)) 334. TBC 334 may be disposed ontometallic layer 332. In this regard, TBC 334 may be applied subsequent tometallic layer 332 being applied to shell 310. Thermal coating 330 maybe applied using plasma spray deposition methods, electron beam vapordeposition, or any other suitable methods. In various embodiments,metallic layer 332 may comprise a nickel alloy. In various embodiments,TBC 334 may comprise a ceramic topcoat. In various embodiments, TBC 334may comprise yttria-stabilized zirconia (YSZ). In various embodiments,TBC 334 may comprise mullite, alumina, Ceria, or any oxide having athermal conductivity which is less than shell 310. In variousembodiments, TBC 334 is corrosion and oxidation resistant.

With reference to FIG. 3B, panel 320 may include one or more panel studs322 for coupling panel 320 to shell 310. Panel stud 322 may extend frompanel 320 and through shell 310. Panel stud 322 may be coupled to shell310, with a fastener such as a nut for example. thermal coating 330 maysurround panel stud 322. Stated differently, panel stud 322 may extendthrough thermal coating 330. In various embodiments, thermal coating 330may be provided locally around panel stud 322 to resist corrosion ofpanel stud 322. In various embodiments, thermal coating 330 may beprovided locally around panel stud 322 to resist corrosion of shell 310at panel stud 322. Thermal protection of panel studs 322 may reduce thelikelihood of a panel stud 322, and associated fastening members such asa washer and/or a nut for example, from being decoupled from shell 310.

With reference to FIG. 3C, dual wall liner 300 may further include aplurality of holes 342 extending through shell 310 and thermal coating330. Plurality of holes 342 may be configures to direct a flow ofcooling air 390 from second side 314 of shell 310 to panel 320. Thiscooling air 390 may then travel through plurality of holes 344 disposedin panel 320 and into hot gas path 392. Hot gas path 392 may be definedby panel 320. In this regard, flow of cooling air 390 may transfer heatfrom shell 310 and panel 320 into hot gas path 392.

In various embodiments, dual wall liner 300 may further include one ormore heat transfer augmentation features, such as trip strip 316 and/orcooling fin 318. Trip strip 316 may extend from second side 314. Tripstrip 316 may be configured to increase a surface area of second side314. Trip strip 316 may be configured to disturb the flow of cooling air390 flowing across second side 314. Cooling fin 318 may extend fromsecond side 314. Cooling fin 318 may be configured to increase a surfacearea of second side 314. Cooling fin 318 may be configured to disturbthe flow of cooling air 390 flowing across second side 314.

With combined reference to FIG. 3A through FIG. 3C, in variousembodiments, an oxidation resistance of metallic layer 332 should begreater than that of the shell material 311 to prevent oxidation and/orcorrosion of metallic layer 332. In various embodiments, metallic layer332 may comprise a material that is at least twice as resistant tooxidation than the shell material 311.

In various embodiments, the thermal conductivity of TBC 334 should belower than that of shell material 311 to prevent heat transfer from hotgas path 392 to shell 310. In various embodiments, the thermalconductivity of TBC 334 may be less than half of that of shell material311.

In various embodiments, although shell 310 may not be exposed to hot gaspath 392 during normal operation, thermal coating 330 may provide addedthermal protection to shell 310 in the event that panel 320 corrodes andholes are formed into panel 320. Stated differently, holes may form inpanel 320 in response to panel 320 corroding due to thermal loadingcaused by hot gas path 392, which may expose shell 310 to hot gas path392. Thus, thermal coating 330 may extend the life of shell 310.

With reference to FIG. 4A, a dual wall liner 400 is illustrated, inaccordance with various embodiments. Dual wall liner 400 may be similarto dual wall liner 300 of FIG. 3A through FIG. 3C, except that dual wallliner 400 includes a thermal coating 430 comprised of metallic layer332. Stated differently, thermal coating 430 is devoid of TBC 334, withmomentary reference to FIG. 3A through FIG. 3C. Thermal coating 430 maythermally protect shell 310 from heat being transferred from got gasesin hot gas path 392 to shell 310. Thus, it is contemplated herein thatthermal coating 430 may be comprised of metallic layer 332.

With reference to FIG. 5, a method 500 for manufacturing a dual wallliner is provided, in accordance with various embodiments. Method 500includes disposing a first layer of a thermal coating over a first sideof a shell (step 510). Method 500 includes disposing a second layer of athermal coating over the first layer (step 520). Method 500 includescoupling a panel to the shell (step 530).

With combined reference to FIG. 4A and FIG. 5, step 510 may includedisposing metallic layer 332 onto first side 312 of shell 310 (step510). With combined reference to FIG. 3A and FIG. 5, step 520 mayinclude disposing TBC 334 onto metallic layer 332 (step 520). Withcombined reference to FIG. 3A and FIG. 5, step 530 may include couplingpanel 320 to shell 310

Benefits, other advantages, and solutions to problems have beendescribed herein with regard to specific embodiments. Furthermore, theconnecting lines shown in the various figures contained herein areintended to represent exemplary functional relationships and/or physicalcouplings between the various elements. It should be noted that manyalternative or additional functional relationships or physicalconnections may be present in a practical system. However, the benefits,advantages, solutions to problems, and any elements that may cause anybenefit, advantage, or solution to occur or become more pronounced arenot to be construed as critical, required, or essential features orelements of the inventions. The scope of the inventions is accordinglyto be limited by nothing other than the appended claims, in whichreference to an element in the singular is not intended to mean “one andonly one” unless explicitly so stated, but rather “one or more.”Moreover, where a phrase similar to “at least one of A, B, or C” is usedin the claims, it is intended that the phrase be interpreted to meanthat A alone may be present in an embodiment, B alone may be present inan embodiment, C alone may be present in an embodiment, or that anycombination of the elements A, B and C may be present in a singleembodiment; for example, A and B, A and C, B and C, or A and B and C.Different cross-hatching is used throughout the figures to denotedifferent parts but not necessarily to denote the same or differentmaterials.

Systems, methods and apparatus are provided herein. In the detaileddescription herein, references to “one embodiment”, “an embodiment”,“various embodiments”, etc., indicate that the embodiment described mayinclude a particular feature, structure, or characteristic, but everyembodiment may not necessarily include the particular feature,structure, or characteristic. Moreover, such phrases are not necessarilyreferring to the same embodiment. Further, when a particular feature,structure, or characteristic is described in connection with anembodiment, it is submitted that it is within the knowledge of oneskilled in the art to affect such feature, structure, or characteristicin connection with other embodiments whether or not explicitlydescribed. After reading the description, it will be apparent to oneskilled in the relevant art(s) how to implement the disclosure inalternative embodiments.

Furthermore, no element, component, or method step in the presentdisclosure is intended to be dedicated to the public regardless ofwhether the element, component, or method step is explicitly recited inthe claims. No claim element herein is to be construed under theprovisions of 35 U.S.C. 112(f) unless the element is expressly recitedusing the phrase “means for.” As used herein, the terms “comprises,”“comprising,” or any other variation thereof, are intended to cover anon-exclusive inclusion, such that a process, method, article, orapparatus that comprises a list of elements does not include only thoseelements but may include other elements not expressly listed or inherentto such process, method, article, or apparatus.

1. A dual wall liner for a gas turbine engine, comprising: a shellhaving a first side and a second side; a panel contacting the shell, thepanel at least partially defining a hot gas path through which a hot gasflows, wherein the first side of the shell faces the panel; wherein theshell includes a thermal barrier coating (TBC) disposed on the firstside of the shell.
 2. The dual wall liner of claim 1, wherein the panelis coupled to the shell.
 3. The dual wall liner of claim 2, furthercomprising a metallic layer in contact with the TBC.
 4. The dual wallliner of claim 3, wherein: the shell is comprised of a shell material;an oxidation resistance of the metallic layer is greater than that ofthe shell material; and the metallic layer is at least twice asresistant to oxidation as the shell material.
 5. The dual wall liner ofclaim 4, wherein the TBC is disposed over the metallic layer.
 6. Thedual wall liner of claim 5, wherein the TBC comprises a lower thermalconductivity than that of the shell material.
 7. The dual wall liner ofclaim 6, wherein the thermal conductivity of the TBC is less than halfof that of the shell material.
 8. The dual wall liner of claim 1,further comprising a heat transfer augmentation feature disposed on thesecond side of the shell.
 9. The dual wall liner of claim 8, wherein theheat transfer augmentation feature is configured to at least one ofincrease a surface area of the second side or disturb a flow of coolingair flowing across the second side.
 10. The dual wall liner of claim 2,further comprising a panel stud coupling the panel to the shell, whereinthe TBC surrounds the panel stud.
 11. The dual wall liner of claim 2,further comprising a plurality of holes extending through the shell andthe TBC, wherein the plurality of holes are configured to direct a flowof cooling air from the second side of the shell to the panel.
 12. A gasturbine engine combustor, comprising: a shell having a first side and asecond side; a panel contacting the shell, the panel at least partiallydefining a hot gas path through which a hot gas flows, wherein the firstside of the shell faces the panel; wherein the shell includes a thermalbarrier coating (TBC) disposed on the first side of the shell.
 13. Thegas turbine engine combustor of claim 12, wherein the panel is coupledto the shell.
 14. The gas turbine engine combustor of claim 13, furthercomprising a metallic layer in contact with the TBC.
 15. The gas turbineengine combustor of claim 14, wherein: the shell is comprised of a shellmaterial; the metallic layer is at least twice as resistant to oxidationas the shell material; the TBC comprises a lower thermal conductivitythan that of the shell material; and the TBC is disposed over themetallic layer.
 16. The gas turbine engine combustor of claim 13,further comprising a heat transfer augmentation feature disposed on thesecond side of the shell, wherein the heat transfer augmentation featureis configured to at least one of increase a surface area of the secondside or disturb a flow of cooling air flowing across the second side.17. The gas turbine engine combustor of claim 13, further comprising apanel stud coupling the panel to the shell, wherein the TBC surroundsthe panel stud.
 18. The gas turbine engine combustor of claim 13,further comprising a plurality of holes extending through the shell andthe TBC, wherein the plurality of holes are configured to direct a flowof cooling air from the second side of the shell to the panel.
 19. Amethod for manufacturing a dual wall liner for a gas turbine engine,comprising: disposing a thermal barrier coating (TBC) on a first side ofa shell; and coupling a panel to the shell, wherein the first side facesthe panel.
 20. The method of claim 19, further comprising: disposing ametallic layer on the first side, the TBC disposed over the metalliclayer.